1. Field of the Invention
The present invention relates to a turbine nozzle for a gas turbine engine and more particularly to a turbine nozzle having curved airfoil that describe a parabola line toward the pressure side along the airfoil stacking axis.
2. Description of the Related Art
A conventional or inventive turbine nozzle is employed in a gas turbine engine, which is described first with respect to its brief configuration.
FIG. 1 is a cross-sectional view showing a brief configuration of a gas turbine engine 1.
The gas turbine engine 1 is employed, for example, as a jet engine in an airplane. It is an engine configured to jet a high-temperature high-pressure combustion gas to provide a propelling force or a rotating force.
The gas turbine engine 1 comprises an engine outer cylinder 3 and, inside the engine outer cylinder 3, a hollow engine casing 5 arranged integrally and substantially coaxial with the engine outer cylinder 3 as base. An annular engine flow path 7 is formed in the engine casing 5. An annular by-pass flow path 9 is formed between the engine outer cylinder 3 and the engine casing 5.
At an inner front portion in the engine casing 5 (an upstream portion seen from the direction of gas flow), interposing the engine flow path 7, a front support frame 11A is formed integrally with the engine casing 5. At an inner rear portion in the engine casing 5, interposing the engine flow path 7, a rear support frame 11B is formed integrally with the engine casing 5. The front support frame 11A and the rear support frame 11B support a low-pressure turbine shaft 13 rotatably via bearings. In addition, the front support frame 11A and the rear support frame 11B support a hollow high-pressure turbine shaft 15 rotatably via bearings and coaxially with the low-pressure turbine shaft 13.
At the front end of the low-pressure turbine shaft 13, a fan 17 is provided to send air into the engine flow path 7 and the by-pass flow path 9.
At a location upstream from the engine flow path 7, a low-pressure compressor 19 is provided. The low-pressure compressor 19 is employed to compress air at low pressure and send it to downstream (seen from the direction of gas flow; and the right side in FIG. 1).
The low-pressure compressor 19 includes an annular airfoil support member 21 located downstream from the fan 17 and arranged integrally with the low-pressure turbine shaft 13. In addition, a row of multi-stage moving airfoils for low-pressure compression 23 is provided at the perimeter of the airfoil support member 21 along the engine flow path 7. Finally, a row of multi-stage stationary airfoils for low-pressure compression 25 is provided inside the engine casing 5 along the engine flow path 7 so as to alternatively interleave with the row of multi-stage moving airfoils for low-pressure compression 23.
At a location downstream from the low-pressure compressor 19 in the engine flow path 7, a high-pressure compressor 27 is provided. The high-pressure compressor 27 is employed to high-pressure compress the air, which has been low-pressure compressed by the low-pressure compressor 19, and send it to downstream.
The high-pressure compressor 27 includes a row of multi-stage moving airfoils for high-pressure compression 29 provided along the engine flow path 7 on the high-pressure turbine shaft 15. In addition, a row of multi-stage stationary airfoils for high-pressure compression 31 is provided inside the engine casing 5 along the engine flow path 7 so as to alternatively interleave with the row of multi-stage moving airfoils for low-pressure compression 29.
At a location downstream from the high-pressure compressor 27 in the engine flow path 7, an annular combustion chamber 33 is formed. The combustion chamber 33 is employed to burn a fuel in a compressed air to produce a high-temperature high-pressure combustion gas.
At a location downstream from the combustion chamber 33 in the engine flow path 7, a high-pressure turbine 35 is provided. The high-pressure turbine 35 is configured to rotationally drive the high-pressure turbine shaft 15 in response to a rotating force caused by expansion of the high-temperature high-pressure combustion gas from the combustion chamber 33.
The high-pressure turbine 35 includes a row of multi-stage moving airfoils for high-pressure turbine 37 arranged on the high-pressure turbine shaft 15 along the engine flow path 7 and operative to rotate in response to the high-temperature high-pressure combustion gas. In addition, a row of multi-stage stationary airfoils for high-pressure turbine 39 is provided inside the engine casing 5 along the engine flow path 7 so as to alternatively interleave with the row of multi-stage moving airfoils for high-pressure turbine 37.
At a location downstream from the high-pressure turbine 35 in the engine flow path 7, a low-pressure turbine 41 is provided. The low-pressure turbine 41 is configured to rotationally drive the low-pressure turbine shaft 13 in response to a rotating force caused by expansion of the high-temperature high-pressure combustion gas from the combustion chamber 33.
The low-pressure turbine 41 includes a row of multi-stage moving airfoils for low-pressure turbine 43 arranged on the low-pressure turbine shaft 13 along the engine flow path 7 and operative to rotate in response to the high-temperature high-pressure combustion gas. In addition, a row of multi-stage stationary airfoils for low-pressure turbine 45 is provided inside the engine casing 5 along the engine flow path 7 so as to alternatively interleave with the row of multi-stage moving airfoils for low-pressure turbine 43. A turbine nozzle in the art or a turbine nozzle according to the present invention includes the row of stationary airfoils for low-pressure turbine 45.
The turbine nozzle is formed from nozzle airfoils (stator vane; hereinafter also referred to as simply “vane”) arrayed in annular. In further detail, the turbine nozzle is formed from a plurality of stator vane, an annular inner band and outer band. For convenience of assembling and disassembling on maintenance, the turbine nozzle as a plurality of nozzle segment generally, may also be referred to as the turbine nozzle in the specification of the present application.
There are two methods to improve the stability to flutter for the vane. One is to increase the natural frequency on the lower order mode of airfoil. The other is to locate the torsional center (torsion center) of lower order modes for airfoil to a stable position.
Generally it is required to change airfoil shape, such a thickness, chord length etc, in order to implement the above methods.
Changing the sectional shape of the airfoil, may have a large influence on aerodynamic performance. In recent design of a high-efficiency turbine. (a low margin (design margin) is remained to change the sectional shape of the airfoil.) Therefore, it may be required to redesign the aerodynamic design in order to change airfoil shape, it is also required to satisfy the mechanical requirements (for example, a mechanical strength) simultaneously. Accordingly, it is difficult in some cases to change the sectional shape of the airfoil.
As disclosed in a patent publication 1 (Japanese Patent Application Laid-open No. H10-196303), there has been known a configuration with stacking of airfoils instead of changing the sectional shape of the airfoil. In this case, the middle portion along the stacking axis of the airfoil is protruded toward the pressure side of the airfoil to curve the airfoil. The stacking is herein defined as to stack the sectional shapes along the stacking axis to construct the shape of the airfoil.
The airfoil disclosed in the patent publication 1 is not provided with indication of specific features of the curve (for example, the shape of the curve and the degree of the curve). Therefore, some feature of the curve can not improve the stability to flutter for vane sufficiently during operation of the gas turbine engine 1 and may make it difficult to produce the turbine nozzle.